AIAA 2010-7015
46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
25 - 28 July 2010, Nashville, TN
Main Belt Asteroid Sample Return Mission Design
John W. Dankanich1
Gray Research Inc., Cleveland, OH, 44135
Damon Landau2
Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 91109
Michael C. Martini3
QinetiQ Group PLC, Cleveland, OH, 44135
Steven R. Oleson4
Glenn Research Center, Cleveland, OH, 44135
and
Andy Rivkin5
Applied Physics Laboratory, Laurel, MD, 20723
The Dawn spacecraft is on its way to rendezvous with two main belt asteroids, Ceres and
Vesta. The science community is already anticipating compelling results that would dictate a
sample return mission. Asteroid sample return missions are already highlighted in the
Decadal Survey and are directly solicited through the New Frontiers mission announcement
of opportunities. There have been numerous studies of near-Earth asteroid sample return
missions. NASA’s In-Space Propulsion Technology project initiated a study to evaluate the
feasibility and propulsion system requirements of main belt asteroid sample return missions.
The mission design trades, results, and propulsion system requirements are presented.
T
I. Introduction
he planetary science division of the NASA science mission directorate ranks missions in terms of science return
as a flyby, rendezvous, land, rove, and ultimately perform a sample return. Unfortunately the missions can
become exponentially more challenging with respect to spacecraft performance capability. The medium class
mission category, New Frontiers, currently specifically calls for missions of interest. One of these mission options is
the asteroid sample return. An asteroid sample return can range significantly from a near-Earth asteroid with a small
gravity well to a very distance asteroid with its own significant gravity. Some asteroids are no more propulsively
challenging than going to the moon, with others cannot be reached within our current propulsion capabilities.
Several institutions have looked at and even proposed relatively simple asteroid sample return missions, and JAXA
returned a sample return capsule from Itokawa on June13, 2010; the world is waiting to see if asteroid materials are
present.
Sample return from near-earth asteroids are quite achievable with today’s technologies using either conventional
chemical propulsion or electric propulsion. NASA concluded that a multi-sample return mission may even be
possible with a single electric propulsion spacecraft.1 While there have been few documented studies to look
beyond small near-Earth asteroids, there is a bounty of science to be found in the main asteroid belt. The Dawn
spacecraft is current on its way to rendezvous with Vesta and then Ceres. Since the formulation of Dawn, Ceres has
1
Lead Systems Engineer, NASA’s In-Space Propulsion Technology Project, M/S 77-4, AIAA Senior Member.
Engineer, Trajectory Design and Navigation, M/S 301-121, AIAA Member.
3
Aerospace Engineer, Systems Analysis Branch, M/S 500-AOS, AIAA Senior Member.
4
COMPASS Lead, Systems Analysis Branch, MS 105-3, AIAA Senior Member.
5
Senior Research Staff, 11100 Johns Hopkins Road.
1
American Institute of Aeronautics and Astronautics
2
Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Go
been upgraded to a dwarf planet. It is expected that the Dawn mission with answer fundamentally science questions,
but also compel the science community to literally “dig deeper” and advocate for a sample return mission.
NASA’s In-Space Propulsion Technology (ISPT) project is tasked with developing propulsion technology to
enhance or enable planetary science missions by increasing mission performance while reducing trip times, cost, and
risk. ISPT products include Aerocapture, the Advanced Materials Bipropellant Rocket (AMBR) engine, NASA’s
Evolutionary Xenon Thruster (NEXT), and smaller investments in the High Voltage Hall Accelerator, and life
testing of the BPT-4000. The AMBR engine, NEXT, and HiVHAC show mission capture for a wide range of
challenging SMD missions. The ISPT project initiated a study to evaluate if a main belt sample return can be
completed using existing technology, one of ISPT’s products, or would require new propulsion technologies. The
objective was to evaluate a representative main belt target and a challenging target. The results of that study are
presented.
II. Science Justification and Payload
In order to ensure a viable mission, the spacecraft design must have a representative science target and also a
representative science payload to determine the required propulsion system performance. In order to justify travel to
the main asteroid belt for a sample return, the target must provide unique scientific insights unavailable in a visit to a
near-Earth object (NEO). This can be due to considerations of geological, compositional, or geophysical diversity.
The largest NEOs are only 30 km in their greatest extent, while main belt objects can be 5-10 times larger (with the
largest one roughly 900 km in diameter). The NEO population contains diverse spectral types, but not in the
proportions found in the main belt due to the specifics of the delivery mechanisms. NEO orbits can pass close to the
Sun, and particular science goals might require targets known to have spent all of solar system history below the
freezing point of water, not always a guarantee for NEOs. We selected two targets based on different science
rationales for a main belt sample return: An easier object rich in calcium aluminum rich inclusions (CAIs) and the
more challenging asteroid/dwarf planet 1 Ceres.
A. 234 Barbara
In the first case, the target asteroid is one of a group that is relatively rare in the main belt, and not known to be
represented in the NEO population at all: objects seen to be rich in calcium aluminum rich inclusions. CAIs are the
oldest solid material in the solar system. They are thus relics of the earliest times in solar nebula history, retaining
traces of nebular properties not seen in other materials. Their sizes as seen in carbonaceous chondrite meteorites are
typically mm to cm.
The CAI-rich asteroids fall into three orbital groups: the Henan and Watsonia dynamical families, and the nonfamily object 234 Barbara. These are all included in the L asteroid spectral class, though not all L asteroids are
known (nor thought) to be CAI-rich.
Science questions a sample return from a CAI-rich asteroid would address include:
1. How did these objects retain/acquire such a large fraction of CAIs?
2. What is the thermal history of CAI-rich objects?
3. How does the non-CAI component of their surfaces compare to known meteorites?
We currently have no way of addressing these questions absent a sample return. With no identified CAI-rich
objects in the NEO population, but multiple objects present in the main asteroid belt, these targets are quite
promising for sample return.
The specific target chosen for the CAI mission study is 234 Barbara. It is one of the objects specified in Sunshine
et al. (2008)2 as a CAI-rich object, and is in a relatively accessible orbit. It also provides additional interesting
science: It has recently been discovered to be a binary asteroid composed of two objects 37 and 21 km in diameter.3
Further detailed information about the system is not currently published, and so was calculated for this study as
needed and as possible. The two components of the Barbara system are calculated to be separated by roughly 100
km, with a mutual orbital period of 24.5 h.
B. Ceres
Ceres is the largest object between Mars and Jupiter. With a polar diameter of ~900 km and an equatorial
diameter of nearly 1000 km, it is roughly twice as large as the next largest main belt objects (2 Pallas and 4 Vesta).
It is also the most massive object between Mars and Jupiter, though it has a surprisingly low density of ~2.1 g/cm3.
Its mass and shape (see below) are sufficient to demonstrate it is in hydrostatic equilibrium, which places it in the
“dwarf planet” category in the IAU classification scheme. It is the only dwarf planet that is not found in the
transneptunian region.
2
American Institute of Aeronautics and Astronautics
Ceres represents the final stage of protoplanets before the planets reached their final form. Its properties, as we
understand them, make Ceres a unique, intermediate object between the rocky terrestrial planets and the icy
satellites of the outer solar system. Products of aqueous alteration like carbonates and brucite (magnesium
hydroxide) have been detected on its surface. Study of a returned sample from Ceres would provide an
understanding of its starting materials, its formation location, and insight into why it is so ice rich in comparison to
its neighbors in the asteroid belt.
Returning a sample from Ceres would allow the following science questions, among others, to be addressed:
1. Is the surface of Ceres part of a primordial crust, a lag deposit above shallow ice, or deeper material
emplaced via cryovolcanism?
2. What relationship does Ceres have to the known meteorites?
3. What are the important dates in Ceres’ evolution?
4. Is there geological evidence for an ancient magnetic field on Ceres?
C. Science Objectives
Science objectives for this study are directly tied to the New Frontiers program Asteroid Sample Return
requirements.
1) "Return a sample to Earth in amount sufficient for molecular (or organic) and mineralogical analyses,
including documentation of possible sources of contamination throughout the collection, return and curation phases
of the mission."
CAI Mission: Meteoriticists have a general consensus that roughly 10 g of material is necessary to get a
representative amount of material for analysis (ref--Sears, I'll dig that up). With CAIs representing 20-40% of the
asteroidal surface, 50 g of sample needs to be collected to obtain 10 g of CAIs. Archiving requirements have varied
with time, but many projects have tried to leave 75%-90% of their sample unanalyzed beyond initial curation, so
that future generations of researchers, equipment, and analysis techniques can have a store of pristine sample. If
75% is archived, an implied total sample mass requirement of roughly 200 g is the result.
Ceres: The sample size for Ceres is not driven by CAI abundance, but the overall reasoning is similar. To obtain
10 g for use in destructive analyses and archive 75%-90% of the total sample implies a sample size of ~50-100 g.
For ease of study, and because the difference was not a mission driver, we assumed the same sample size for the
Ceres mission as the CAI mission.
2) "Map the surface texture, spectral properties (e.g., color, albedo) and geochemistry of the surface of an
asteroid at sufficient spatial resolution to resolve geological features (e.g., craters, fractures lithologic units)
necessary to decipher the geologic history of the asteroid and provide context for returned samples."
CAI Mission: Prior to sampling, a reconnaissance phase will be required. The mission in the study is equipped
with visible and near-IR imaging and spectral capability and a LIDAR in order to identify the best spot to sample
and understand its relationship to the rest of the target. A map of the surface to 3 m/pixel will identify any regions
like the ponds on Eros and Itokawa, which would be easier areas to obtain a sample, though provision for validation
of sampling sites via targeted higher-resolution imaging is also present in the timeline. A 3 m/pixel map will also
provide a full accounting of the geologic diversity present on the target.
With the camera chosen (an analog of MDIS on MESSENGER), multispectral imaging at 20 m/pixel with a
wide-angle camera (WAC) can be obtained while at the same target distance as the narrow-angle camera (NAC)
obtaining 3 m/pixel imagery. Infrared coverage is provided out to 3.5 µm by including a spot spectrometer like the
MASCS instrument on MESSENGER, which can look at spots ~100 m in size from the mapping orbit. The sum of
the spectral coverage will provide compositional context for the returned sample as well as constrain any
heterogeneity and guide the choice of sampling site.
Ceres: The Dawn mission will obtain reconnaissance data, obviating the need for that phase of a Ceres sample
return. It is assumed here that Dawn will provide 75 m/pixel imaging of Ceres’ entire surface and a relatively small
number of possible sampling sites will be identified prior to launch. These areas will be characterized in detail using
the optical instruments in order to provide the detail for geologic context (~3 m/pixel) and geochemistry (~10-20
m/pixel) to fill this requirement.
Because several instruments were descoped from the Dawn mission, including a LIDAR and a magnetometer,
their inclusion in a Ceres sample return would provide the opportunity for bonus science at potentially small
additional cost. Inclusion of such instruments would require an operations phase, but the time spent at Ceres is
driven by the dynamics required for a return to Earth, and may be something the mission can accommodate.
However, a magnetometer and LIDAR are not key instruments for this sample return, and could be descoped easily.
3) "Document the regolith at the sampling site in situ with emphasis on, e.g., lateral and vertical textural,
mineralogical and geochemical heterogeneity at scales down to the sub-millimeter."
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American Institute of Aeronautics and Astronautics
CAI mission: The remote sensing instruments used for the previous objective will also be used as appropriate and
possible to achieve this objective. Observations will be performed upon ascent and decent (and on the surface?) in
order to provide before and after looks at the sampling site to determine the extent of space weathering and as a
further measure of context. These investigations can both use observations of disturbed vs. undisturbed soil, and
study of how spectral properites change with depth in the sampled area.
Ceres mission: The sampling on Ceres occurs during a landed phase rather than a touch and go. The required
imaging and documentation will be achieved using microscopic imagers and panoramic cameras similar to what is
used on the MER mission. In addition, ascent and descent imagery will take advantage of the WAC/NAC in order to
provide larger-scale context. In addition, small regolith probes on the landing pads are included, similar to what is
carried by the Rosetta lander, in order to measure the conductivity and permittivity of Ceres’ regolith.
D. Payload
The science goals are used to derive science objectives. From the science objectives, a set to desired data with
measurement requirements can be formulated. The measurement objectives then lead to a selection of a mission
specific instrument sweet ultimately required to meet the over arching science goals. The instrument traceability
matrix is provided in Table 1.
Table 1. Science traceability matrix for main belt sample return missions.
The full mission payload consists of optimized instruments designed to support both orbital and landed
operations. Feature mapping from orbit will be done with wide- and narrow-angle imagers, a laser altimeter, and an
infrared spectrometer. A permittivity probe for characterizing the dielectric properties of the regolith is included on
the Ceres mission. The instruments work to help identify promising sampling sites as well as achieving important
science on their own. The CAI mission and Ceres mission carry different instruments in their baseline due to the
opportunity to build upon earlier missions for Ceres, though at the floor they are the same. The instrument specifics
can be found in the appendix, but the summary table is shown in Table 2.
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Table 2. Representative science payload for the two sample return missions. Yellow is the baseline payload
and green is the floor science payload.
III. Spacecraft Design and Requirements
The spacecraft design is an iterative process with the propulsion system influencing the other subsystems. The
COMPASS team iterated with the detailed subsystem elements for a full spacecraft design. Based on the propulsion
system performance a spacecraft mass equipment list (MEL) is developed. The COMPASS team uses the AIAA
standard for mass growth, contingency, and margin policy.4 Growth factors are applied to each subsystem, after
which a total system growth of the design is calculated. The COMPASS design standard operation is to carry
additional system level growth of 30% on the dry mass. Wet mass growth is carried at either the propellant mass
level, such as for propellant residuals, or through the inclusion of ΔV margin. The MEL is critical for viable mission
trades to ensure that the propulsion system and system implications are properly captured. The detailed point design
MELs and spacecraft configurations are provides in session IV. Figure 1 is a spacecraft configuration for a small
asteroid sample return spacecraft designed for Barbara, shown without the solar array, and a large asteroid sample
return spacecraft designed for the size and gravity of Ceres.
Figure 1. Preliminary spacecraft design for Barbara (left) and Ceres (right).
A. Mission Design Trades
The intent of the study is to evaluate state-of-the-art propulsion system performance and determine if any new
propulsion technologies are required to enable the missions within the New Frontiers cost cap. The only limitation
imposed was to stay within the medium launch vehicle capabilities. For this purpose, the highest performance
launch vehicle considered is the Atlas 551.
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American Institute of Aeronautics and Astronautics
B. Ballistic Solutions
In general, if a mission can be completed using a chemical propulsion system, the mission cost is expected to be
lower. Also, landing with large solar arrays associated with solar electric propulsion systems can add mission risk
and complexity. The first option was to evaluate ballistic solutions for the use of commercial thrusters or the AMBR
engine. A broad search of potential ballistic sample return solutions was conducted using STAR, a Lambert based
solver for rapid assessment of non-optimal ballistic solutions. The search included 211 main belt targets with
diameters greater than 100km. The search was limited to a maximum trip time of 12 years with launch dates from
2018 – 2023.
A broad search was completed for various combinations of Mars fly-bys both before and after the asteroid
encounter. The non-optimized solutions are shown in Fig. 2. One of the highest performance targets was chosen for
optimization to provide the best performance capability of an asteroid sample return using a chemical propulsion
system. The optimized solution to Flora is shown if Fig. 3. The results show that a spacecraft launched to a C3 of
18.5 km2/s2 can perform two Mars gravity assists and arrive at Flora four years after launch with a capture ΔV
requirement of 1.46 km/s and then depart Flora with a ΔV requirement of 1.73 km/s to perform two Mars gravity
a)
b)
Figure 2. Unoptimized ballistic main belt asteroid sample return trajectories using EMMaMME a) (left) and
EMMaME b) (right) flyby sequences.
assists on the way back and arrive
with an entry velocity of 14.4 km/s for
a total mission time just under 11
years. This is a total ΔV of 3.2km/w,
which might be on the limits of a
chemical propulsion post launch ΔV.
On at atlas 551, before margins, this
could deliver approximately 1500 kg
back to Earth. Flora serves as a test
case that a chemical sample return
may be possible from a small and easy
target, but it would be propulsively
challenging and the largest chemical
post launch ΔV ever performed by a
chemical system on a NASA science
mission. Figure 4 provides the
optimized Ceres solution with a ΔV
over 6.5 km/s. Based on these results
it is not practical to persue a chemical
propulsion based main belt sample
return mission within the New
Frontiers mission category.
Figure 3. Optimized ballistic solution to Flora.
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C. Low Thrust Solutions
Low-thrust mission trades and point
designs were performed using the
Mission
Analysis
Low-Thrust
Optimization
(MALTO)
tool.5
Preliminary low-thrust trajectories have
been evaluated for inner, regular, and
outer main belt targets. The vast
majority of main belt asteroids range in
semimajor axis from approximately 2.1
– 3.3AU. The initial trades included
evaluating the BPT-4000 Hall thruster,
the HiVHAC, and NEXT. Trades
included trip time evaluation, solar
array
power
requirements,
and
propulsion
system
configuration.
Figure 5 shows notional solutions for
Tea, Ceres, and Paracelsus located at
2.2, 2.8, and 3.2 AU respectively.
Figure 4. Optimized ballistic solution to Ceres.
a)
b)
c)
Figure 5. Notional solutions to a) Tea, b) Ceres, and c) Paracelsus.
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Because the ΔVs for a main belt sample
return mission are expected to be very high,
Hall thrusters were not anticipated to have the
highest performance. The NEXT thruster
outperformed both the BPT-4000 thruster and
the HiVHAC for even for just the outbound leg
to Vesta for all power levels evaluated. Figure 6
illustrates the performance to Vesta for the
three propulsion systems as a function of solar
array power with a constrained launch mass of
2000kg using an Atlas V 401. This does not
preclude the use of Hall thrusters for this
mission, but only that the NEXT IPS is
expected to be the highest performing system
for a main belt asteroid sample return. Based on
this finding, NEXT is the baseline propulsion
system for the detailed point design. If none of
Figure 6. Delivered mass to Vesta for the BPT-4000,
the current options provided a viable solution,
HiVHAC, and NEXT options.
trades for optimal Isp, efficiency trades, lifetime
capabilities, etc. would have followed to quantify the required system performance to enable these missions.
D. Ceres Descent and Ascent Propulsion
For small asteroids, the descent and ascent propulsion can be small commercial off-the-shelf (COTS) thrusters.
However, the Ceres mission challenged the performance capabilities of SOA COTS systems. Because of the large
gravity well of Ceres, the low-thrust propulsion system will bring the spacecraft down to a very low orbit before
decent. The chemical propulsion system will perform the decent and ascent placing the spacecraft into a 10 km x
100 km elliptical orbit. The electric propulsion system will then raise the orbit to escape and complete the transfer
back to Earth.
The ascent requirements were calculated using the Optimal Trajectories by Implicit Simulation 4 (OTIS4)
program. Trajectories from the surface of Ceres to a wide range of orbits were evaluated. Trades assumed a Ceres
landed mass of 1500 kg; various thrust levels and specific impulse trades were made based on potential propulsion
system configuration. The launch was due east from the equator to the final orbit. A 10 km x 100 km orbit was
chosen for the relatively low chemical ΔV of 425 m/s and to maximize the benefit of the SEP system. The decent ΔV
is an approximation by assuming 1.11 times the ascent requirement. This assumption is based on observed ratios
from previous mission, e.g. Apollo decent and
ascent. After the orbit selection, and ΔV
calculations, the propulsion system performance
iterated through the COMPASS design process
for a suitable thruster level and corresponding Isp
to a commercially available thruster. The mission
converged to a landed mass at Ceres of nearly
2,150 kg corresponding to a weight of 142 lbf on
the surface of Ceres and the design requirement
of a 200 lbf total propulsion system thrust level.
Rather than deep throttling of a single engine, a
set if 8 x 25 lbf Marquardt R-1E thrusters were
selected with an assumed Isp of 285 seconds. The
total spacecraft wet mass savings is on the order
of 3 kg for each second of Isp improvement. Even
though the mission did converge, the
ascent/decent engines were identified as a source
for reasonable performance gain through a small
investment to qualify a higher performance midrange thrust, ~25-50 lbf, bipropellant thruster.
Similar improvements have been made and
Figure 7. Ceres ascent to a 10 km x 100 km orbit.
qualified for 100-lbf thrusters.
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IV. Final Mission Design and
Spacecraft Configuration
Using the COMPASS missions
design team, the iterative process led to
mission point designs to add credence to
the propulsion system requirements.
Detailed spacecraft design reports are
available for both Barbara6 and Ceres7.
For reference, the final cost of the
Barbara mission was approximately
$400M; insufficient margin for a
Discovery mission, but it fits easily
within New Frontiers with more than
50% cost margin. The final cost of the
Ceres mission was estimated at
approximately $517M, which is close,
but still fits within the New Frontiers cap
with the required 25% margin.
A. Barbara
1. Detailed Mission Design
Figure 8. Baseline trajectory for Barbara.
The baseline point design for the
Barbara sample return mission is shown in Fig. 8. The baseline trajectory launches to a relatively high C3 at 38.8
km2/s2 on August 16, 2021 and arrives at Barbara in October of 2025. The spacecraft then stays in the vicinity of
Barbara for the orbital and landed phases of the mission for 180 days before departing back to Earth. The return leg
takes another 3.4 years and arrives at Earth with a constrained V∞ of 7km/s. The total mission duration from Earthto-Earth is 8 years. The constrained V∞ provides an entry velocity of approximately 13.1 km/s. The baseline solution
performs a total post launch ΔV of 12.85 km/s while consuming 407.43 kg of Xenon and operating for 2,279 days
for a total impulse of nearly 1.4 x 107 N-s. An additional 35 kg of propellant is carried from trajectory corrections,
propellant residuals, margin, etc. The entire mission is designed for a 1+1 NEXT system with a single operating
thruster planned for the entire mission and a cold spare string carried for redundancy.
The baseline mission is designed with 8.5 kW of solar power at 1 AU at the end of life and with 276 W of
housekeeping power. The propulsion system requirements include an additional xenon load of 35kg for navigation
and trajectory corrections, propellant residuals, etc. for a total loaded mass of 442.7kg. The spacecraft can fit within
the Atlas V 401 launch vehicle; with a launch capability of 1,500 kg to the required C3.
The main belt sample return trajectory does
not have significant cost periods, which can be
a mission risk. A more robust trajectory with
larger coast periods may be possible with a
higher thrust system or an increase in available
power. Figure 9 shows solar array power profile
and the PPU input power over the course of the
mission. The thruster has a short coast period
near launch and then thruster continuously until
920 days, then there is a 120 day coast period
before the thrusters must operate continuously
until the arrival at Barbara. The mission is
modeled with a 90% duty cycle to allow for
missed thruster periods for communications and
unplanned outages.
The spacecraft departs the Earth near 1 AU
and never comes closer to the sun until
returning back to Earth at the end of the
Figure 9. Solar array and PPU input power for the duration of mission. The maximum distance from the Earth
the Barbara sample return mission.
and the spacecraft is 4 AU, and this occurs
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during the science phase of the mission. This
means that the Earth and the spacecraft are in
opposition with the Sun as the spacecraft is
maneuvering, landing, and performing the
sample collection. However, because of the
inclination of the target, there will not be any
significant
communication
loss
from
occultation. The spacecraft distances from the
sun and Earth are shown in Fig. 10.
2. Top Level Mass Equipment List (MEL)
Through iteration with the mission design,
the COMPASS team arrived at a closed
spacecraft design to meet all of the mission
objectives and carry the entire science
payload with the required system level
margins. The MEL for the Barbara mission is
provided in Table 3. The mission has a total
launched wet mass with growth just under
1,300kg including 55kg for the science
payload.
Figure 10 Spacecraft distances from the sun, Earth, and
3. Spacecraft Configuration
In addition to the MEL, the spacecraft configuration is necessary to ensure the placement of the thrusters,
sensors, solar arrays, science instruments, etc. can allow for a low risk implementation of the mission. The baseline
configuration for the Barbara main belt sample return mission is shown in Fig. 11.
Table 3. Top level mass equipment list for the Barbara sample return mission.
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Figure 11.
Baseline configuration for the Barbara main belt asteroid sample return spacecraft.
B. Ceres
1. Detailed Mission Design
The sample return mission from Ceres is significantly more challenging. The target is more challenging partially
due to the farther distance that must be traveling to reach Ceres, but primarily because of the large propulsion
system that must be used to land and depart from the surface of Ceres with its relatively large gravity well. Because
of the additional mass requirements, a single thruster cannot push enough mass to Ceres and stay within the total
mission time of 12 years. For the baseline design, the propulsion system is based on a 2+1 configuration, with two
concurrently operating thrusters for higher thruster and one cold spare string for redundancy.
The baseline trajectory for the sample return mission from Ceres is shown in figure 12. This baseline trajectory
launches with a C3 of 32 km2/s2 on December 20, 2020 and arrives 5.4 years later in May of 2026. The spacecraft
then stays in the vicinity of Ceres for orbiting and landing science and sample collection for 240 days before
departing back to Earth. The spacecraft must spiral down to a low Ceres orbit to reduce the chemical landing
propulsion requirement and also spiral to escape. The spiraling time inbound and to escape are 104 and 57 days
respectively. This spiraling time, in addition to science, sampling, and margin requires the spacecraft to stay with
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Ceres vicinity for 240 days before
departure. The return leg takes an
additional 3.5 years and arrives at earth
with a V∞ of 6.98 km/s. The total mission
duration from Earth-to-Earth is 9.5 years.
The constrained V∞ provides an entry
velocity of approximately 13.1 km/s. The
baseline solution performs a total post
launch ΔV of 17.5 km/s while consuming
916.06 kg of Xenon and operating for
2,700 days for a total impulse of more
than 3.6 x 107 N-s.
The baseline mission is designed with
20 kW of solar power at 1 AU at the end
of life and with 300 W of housekeeping
power.
The
propulsion
system
requirements include an additional xenon
load of 89 kg for navigation and trajectory
corrections, propellant residuals, spiraling
to a low Ceres orbit, etc. for a total loaded
Figure 12. Baseline trajectory for Ceres.
mass of 1,005 kg. This also drives the
propellant throughput requirement for
each NEXT thruster to approximately 500kg. The spacecraft fits within the Atlas V 551 launch vehicle at a launch
mass of 3,191.1 kg and a launch capability of 3,545kg to the required C3.
Even with the additional thruster, the delivered mass performance is limited by thrust, with no sizable costing to
Ceres; this will carry mission risk. Coast periods can be introduced by increasing the total mission time with a
moderate propellant penalty. The mission is modeling with an assumed 90% duty cycle and significant power
margin which could alleviate the missed thrust period risk. Figure 13 illustrates the power available out of the solar
array and PPU input power over the transfer times of the mission.
Like the Barbara mission, the Ceres mission launches just below 1AU from the sun and with a maximum
distance of nearly 3AU occurring near the spacecraft arrival date. The spacecraft has a maximum distance from
Earth of 3.9 AU, but the spacecraft is on the same side of the sun during the sampling phase of the mission with an
Earth – Spacecraft distance of 2.78AU at arrival and 2.64AU at departure. Spacecraft distances over time are shown
in Fig. 14.
Figure 13.
Baseline trajectory for Ceres.
Figure 14.
Baseline trajectory for Ceres.
2. Top Level Mass Equipment List (MEL)
The mass of the Ceres sample return mission is significantly higher than the Barbara case. The spacecraft has a
lander stage that will be carry the propulsion system for the landing phase and the ascent system. The lander stage
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will then separate from the vehicle so that the electric propulsion system does not have to push the entire mass back
to Earth. The top level MEL is shown in Table 4. The spacecraft has a total wet mass with growth of 3,158 kg. Of
the wet mass, more than 1,800 kg is propellant with more than half as Xenon for the electric propulsion system.
Table 4. Top level mass equipment list for the Ceres sample return mission.
3. Spacecraft Configuration
The original options for the Ceres mission including landing and returning the entire spacecraft, have a
mother/daughter spacecraft where the mother spacecraft would stay in orbit and perform autonomous rendezvous
and docking and the daughter retrieved a sample, or landing the entire spacecraft and staging before the return to
Earth. Returning the entire spacecraft put too much demand on the electric propulsion system, using a
mother/daughter spacecraft pushed the mission beyond the cost limits of a New Frontiers missions; so landing the
entire spacecraft and staging before departure was the final mission concept selected, but it also has risks. The solar
arrays have additional mass for the landing and thruster loads and staging always adds mission complexity. The final
configuration of the Ceres sample return mission is shown in Fig. 15.
V. Conclusion
In general, primary chemical propulsion cannot be used for a main belt sample return mission. For both missions
selected for evaluation, the primary electric propulsion system is limited by thrust, but the very high ΔV’s do not
trade well with Hall thrusters. Both missions pushed the throughput limits of the NEXT IPS, but are expected to be
within its designed capabilities. Increased power had the largest impact on the propulsion system, but power
recommendations are beyond the scope of the ISPT project. A qualified mid-range, ~25-50 lbf, higher Isp
bipropellant thruster can save significant mass for the larger asteroid sample return missions. During the study, it
was observed that an acceptable method to relieve the requirement of a spare thruster string could lead to significant
savings in mass and cost. Overall, the results show the NEXT ion propulsion system may be sufficient for a wide
range of main belt sample return missions without the need for additional technology investment.
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Figure 15. Baseline configuration for the Ceres main belt asteroid sample return spacecraft.
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Appendix
A. Science Payload Details
The payload works in concert to make detailed complementary measurements and achieve the science objectives
within the mass, power, cost, and operational constraints consistent with a Discovery-class program. The baseline
mission carries all of these instruments. The science floor mission carries a subset of these instruments: several
possible combinations of instruments can meet the floor-level science objectives. For instance, while a camera is
required for any version of this mission, The prospective Principal Investigator will define the specific the science
floor within the constraints of the final Discovery Announcement of Opportunity and select the best combination of
instruments to achieve this floor. Note also that some of this instrument complement may lend itself easily to foreign
contribution, such as the Rosetta-derived surface probes. As much as half of the baseline instrument complement is
likely to be suitable for foreign contribution.
B. Imager
The imager concept for this mission is based on a greatly simplified version of the MESSENGER Dual Imaging
System (MDIS) shown in Fig. A-1, which contains both a wide-angle camera (WAC) and a narrow-angle camera
(NAC) on a common pivot platform. The orbit about a Trojan asteroid is slow enough that all mapping objectives
can be achieved with spacecraft pointing alone, as demonstrated on the Near-Earth Asteroid Rendezvous (NEAR)
mission, thereby eliminating the need for the pivot platform. This simplification is significant because it reduces
instrument complexity, alignment concerns, and on-orbit calibration. The thermal control system, which is the most
complex part of the MESSENGER MDIS instrument, can be greatly simplified by eliminating all of the special
thermal controls for operation at Mercury. The beryllium radiators, heat pipes, phase-transition wax packs, and a
large part of the structure are all removed. These simplifications taken together, and allowing for larger optics (~75
mm), bring the total instrument mass to an estimated 4.0 kg.
The imager focal planes are based on the New Horizons Long Range Reconnaissance Imager (LORRI)
instrument, which employs low-noise, high quantum efficiency CCDs that are better optimized for low-light
conditions and long exposure times. Global mapping is achieved with the WAC, which covers a ~15° field of view
(FOV) and features a 12-position filter wheel spanning a wavelength range from 395 to 1040 nm. In the open filter
position, the WAC can also serve as a descent imager. The NAC a has a 1.5° FOV and can deliver high-resolution,
panchromatic images of targeted surface features. The CCDs themselves are 1024  1024 element arrays with 14µm pixels.
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1024  1024 CCD, 14-µm pixels
Narrow angle camera (NAC)
1.5°  1.5° Field of view
panchromatic
25-µrad IFOV
Wide angle camera (WAC)
12-position filter wheel
395–1040 nm
250-µrad IFOV
Mass 4.0 kg
Power 7.0 W
Data rate 3 Mbps
Figure A-1: MESSENGER Dual Imager. The dual imager eliminates the pivot platform and simplifies the
thermal design to achieve a mass of 4 kg.
C. Laser Altimeter
The laser altimeter is based on a straightforward unit developed for NEAR, shown in figure A-2. The instrument is
configured as a direct-detection laser radar with the transmitter and receiver in a bistatic arrangement. The
transmitter is a diode-pumped Cr:Nd:YAG (1.064-µm) laser. The Q-switched laser emits 8-ns pulses at 15 mJ per
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pulse, with a repetition rate of up to 2 Hz. The separate receiver uses an avalanche-photodiode (APD) detector
coupled to a 7.6-cm clear aperture Dall-Kirkham telescope. A semi-transparent door protects the optics during
integration and launch. The sensitive range is 200 km, with a range accuracy of 1.5 m and a range resolution of 0.3
m. The electronics will be updated to incorporate the APL-developed time-of-flight (TOF) ASIC, include newer
implementations of the power electronics and instrument processor, and replace the 1553 serial data interface with a
low-power LVDS serial interface.
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200-km range
1–8 Hz operation
Laser wavelength 1.064 µm
Range accuracy < 1.5 m
Range resolution < 0.3 m
Mass 5.0 kg
Power 8.0 W
Data rate ~ 50 bps
Figure A-2: The NEAR LIDAR. This mission can make use of heritage laser optics with an updated design of
the electronics to reduce mass and power.
D. Infrared Spectrometer
Given the limited payload resources for an asteroid sample return mission, a complex infrared spectrometer such
as the Compact Reconnaissance Imaging Spectrometer for Mars (CRISM), with its 33-kg mass and 45-W power
dissipation, is not tenable. A simpler, lower-mass and lower-power instrument is needed, with a complexity more
similar to the MESSENGER Mercury Atmospheric and Surface Composition Spectrometer (MASCS) instrument
(Figure A-3). The MASCS instrument concept can be further simplified by eliminating the ultraviolet and visible
portions of the instrument. The wavelength range will be extended out to 4.5 µm by replacing the photomultiplier
tube (PMT) detector with a cooled HgCdTe solid-state detector, which is available in a sealed package equipped
with an integral thermoelectric cooler (TEC). This IR spectrometer has an overall mass of 3 kg and an operating
power dissipation of 11 W.
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Detector 1-mm dia. HgCdTe with integrated
multi-stage TEC
Wavelength range 2.25 to 4.5 µm
Spectral resolution 10 nm
Spatial resolution 0.025°
Mass 3.0 kg
Power 11.0 W
Data rate ~ 300 bps
Figure A-3: MESSENGER MASCS Instrument. The IR spectrometer is similar in complexity to that shown, but
eliminates the UV and visible portions and replaces the PMT detector with a cooled HgCdTe detector.
E. Permittivity Probe
For very little additional resources, each landing foot can be equipped with a simple permittivity probe to
characterize the dielectric properties of the regolith. Figure A-4 shows the probe developed for the ROSETTA
mission. The probe employs an active sensor that can sweep frequencies over the range 10 Hz to 10 kHz, which is
considered optimal for detecting water. Capacitive coupling of the sensor signals makes the probe readings
insensitive to the contact properties between each landing foot and the surface. A small preamplifier is co-located
with each sensor while the supporting electronics can be located with the other instrument electronics. The simple
processing needs of the permittivity probes can likely be absorbed into one of the other landed instruments.
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•
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Active probe measures diurnal variations in
permittivity
Probe in each landing foot
Capacitive coupling insensitive to contact
properties
10 Hz – 10 kHz frequency range optimum for
H2O detection
Mass 1.0 kg
Power 1.0 W
Figure A-4: Sole and cover of a ROSETTA landing foot. The sole (right, ring diameter: 10 cm) contains an
electrode for the permittivity probe. The preamplifier is mounted on the inner side of the lid (left).
F. Arm-mounted camera
While on the surface, an arm-mounted camera is used to achieve the required spatial resolution and
documentation of the sampling site, as well as to help verify the sampling events. The Barbara mission includes the
Athena Microscopic Imager (MI) to achieve these goals. The MI optics used a fixed-focus f/15 design, providing 30µm/pixel resolution and a 31x31 mm field of view. This instrument has been flown on the Mars Exploration Rovers
(MER), and should work well for an asteroid sample return.
G. Descent Imager
In order to ease the acquisition of pre- and post-sampling context imaging, a descent (and ascent) imager is
included on the CAI mission part of the Barbara mission. This will allow the WAC/NAC combo to be optimized for
global mapping. The Mars Descent Imager (MARDI) included on the Mars Polar Lander and Mars Phoenix
spacecraft was included in the mission study. It is a 9-element refractive system, capable of cm-scale imagery over a
66 degree field of view at 10 m altitude.
H. Landed Camera
The Ceres mission includes a landed panoramic camera to provide context for the landed phase. The MER Pancam
was included in the spec to represent this instrument. Pancam can provide multispectral stereo panoramic imagery,
with information about the site’s morphology and the sampling area pre- and post-acquisition. Pancam’s angular
resolution is 0.28 mrad/pixel, with a focus from 1.5 m to infinity, and a field of view of 16.8 x 16.8 degrees.
Acknowledgments
The work described in this paper was funded in whole or in part by the In-Space Propulsion Technology
Program, which is managed by NASA's Science Mission Directorate in Washington, D.C., and implemented by the
In-Space Propulsion Technology Project at the John Glenn Research Center in Cleveland, OH. The program
objective is to develop in-space propulsion technologies that can enable or benefit near and mid-term NASA space
science missions by significantly reducing cost, mass or travel times. The author gratefully acknowledges the
contributions of the entire COMPASS team for the spacecraft design and mission requirements.
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